Bleed for compressor



Feb. 3, 1970 l FgA N, A'BlLD` ET AL 3,493,169

BLEED FOR COMPRESSOR y Filed June s, 1968 United States Patent O 3,493,169 BLEED FOR COMPRESSOR Robert N. Abild, New Britain, and Charles B. Mayer,

Tolland, Conn., assignors to United Aircraft Corporation, East Hartford, Conn., a corporation of Delaware Filed .lune 3, 1968, Ser. No. 733,871 Int. Cl. F04d 25/16, 29/30 U.S. Cl. 230-119 2 Claims ABSTRACT OF THE DISCLOSURE This invention has been reported as a Subject Invention under an Air Force contract.

BACKGROUND OF THE INVENTION In the use of multistage compressors in aircraft engines, for example, it is desirable to obtain air for use other than in the engine or even for use in cooling the turbine. For either purpose an elevated pressure is desirable without, however, having the air heated during the compression process in the last or centrifugal stage. Where the air is turned from an axial direction of flow toward a radial flow through the compressor rotor there is a substantial Velocity pressure that, combined with the static pressure at this point, produces a substantially elevated pressure but at a much lower temperature than prevails at the centrifugal compressor outlet. This total pressure is high enough to be usable for turbine cooling and is at such a temperature as to be effective as a cooling medium.

SUMMARY O-F INVENTION A feature of the invention is the -bleeding olf of air or other gas from the centrifugal last stage of a multistage compressor through axial passages in the centrifugal rotor located at the point where the rotor turns the flow from axial toward radial ow, thereby to develop the total pressure available in the fluid being compressed at this point.

j In accordance with this invention, the centrifugal compressor rotor has axially extending, forwardly open passages therein to receive a part of the axially flowing compressible fluid in the compressor, these passages being located at a point where the rotor begins to impart a radial ilow to the gaseous Huid. The dynamic pressure of the fluid is converted into static pressure within these passages for an effectively higher total pressure. Suitably arranged passages in the rotor direct the gaseous fluid to a plenum chamber adjacent to or surrounded by the rotor.

One particular feature a two-part centrifugal rotor in one part of 'which the axial passages are provided, with intel-communicating radial passages provided in one or both of the mating surfaces of the rotor through which the bleed gas is conducted to a centrally located axial passage in the rotor.

BRIEF DESCRIPTION OF THE DRAWING The single gure is a longitudinal sectional view through the compressor.

DESCRIPTION OF THE PREFERRED EMBODIMENT The compressor shown has a casing 2 with a number of rows of axial flow stator vanes 4 projecting inwardly therefrom and alternating with rows of blades 6 on the rotor 8-. Gaseous fluid entering the inlet 10 is compressed by these alternating blades and vanes and is discharged in an axial direction into the last or centrifugal stage 12. This stage includes a rotor 14 having a curved surface 16 that is substantially axial at the inlet and merges to a radial surface at the outlet. This surface has a plurality of vanes 18 projecting therefrom toward a cooperating curved surface 20 on the casing. From the rotor 14 the fluid being compressed is discharged through a diffuser 22.

The rotor is supported in spaced bearings 24 and 26. The bearing 24 is supported through a series of struts 28 and a support disk and the bearing 26 is located at the inner perphery of a support 32 extending inwardly from the casing at the diffuser 22. A seal ring 34 on the rotor and a cooperating ring 36 on the support 32 define a chamber 38 communicating through ports 40 in the rotor with a second chamber 42 within the rotor.

In accordance with the invention, the centrifugal rotor has axially extending passages 44 located. at a point where the curved surface 16 has begun to approach a radial direction and open forwardly in such a way that a part of the gaseous uid moving in an axial direction from the last row of stator vanes 4 will enter these passages thereby providing at this point a total pressure that is the sum of the static pressure in this part of the compressor and the dynamic or velocity pressure resulting from the axial motion of the fluid.

From the passages 44, intersecting radial passages 46 guide the bleed fluid radially inward to the chamber 42 and thence to chamber 38. From these chambers the bleed fluid is utilized as desired either in being circulated to the turbine which drives the compressor for cooling the vanes or blades of the turbne, or for any other particular use 'for which the gaseous fluid is adapted.

It will be apparent that at this point the centrifugal stage is just beginning its compression function during which the temperature of the fluid is substantially increased as a result of the compressing operation. Thus, the bleed passages above described are located to bleed the fluid at a point where the pressure is relatively high but before the fluid is heated too much by the compressing function of the centrifugal stage.

The centrifugal rotor 14 may be\made in the form of two disks 14a and 14h, the former having the inlet portion of the rotor and the latter having the outlet portion. The radial passages 46 are formed in one or both of the mating surfaces of these disks, being located in the specific arrangement in disk 14b. In this way the passages 46 may terminate at the intersection with passages 44 and need not break through radially into the curved surface 16. Suitable bolts, not shown, may hold the disks together or they may otherwise be secured together after the formation of the passages therein.

We claim:

1. A multistage compressor having at least one axial ilow entry stage defining an annular axial flow path and a final centrifugal stage, said compressor comprising a rotor having at least one row of axial flow blading and a centrifugal rotor having an arcuate surface downstream of the axial blading, a stator having at least one axial stage of stator vanes and an arcuate wall downstream thereof cooperting with the centrifugal rotor, said centrifugal stage changing the flow of gas being compressed from an axial direction to a substantially radial direction, and bleed passages in said centrifugal rotor having the inlet ends thereof extending substantially axially of the rotor and terminating at said arcuate surface at a point where the arcuate wall extends substantially axially and forms an extension of the inner wall of the axial flow path and where the gas flow is substantially axial from the axial stage whereby the gas enters the bleed passage with no significant change in the direction of ow.

2. A multistage compressor as in claim 1 in which the centrifugal rotor is made up of two mating discs in one of which the axially extending bleed passages are formed, and in the mating surfaces of which radial passages are formed to intersect at their outer ends with the axially extending bleed passages.

References Cited UNITED STATES PATENTS 1,151,964 8/1915 Peterson 103-112 1,323,412 12/1919 Schorr 103-112 1,473,802 11/1923 woock er a1. 103-112 8/1932 Schlachter 103-112 5/1942 Birmann 60-59 XR 12/1952 Parducci 253-39.1 7/1954 Craig et al. 253-39.1 XR 5/1955 Wood 253-39.1 6/1955 Birmann 253-39.l XR 9/1963 Rockwell 230-119 XR FOREIGN PATENTS 7/1914 Germany. 2/1931 Germany. 3/1949 Great Britain. 5/ 1928 France. 4/1939 Italy.

HENRY F. RADUAZO, Primary Examiner U.S. Cl. X.R. 

